A gas turbine engine typically includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure compressor includes annular arrays (“rows”) of stationary vanes that direct air entering the engine into downstream, rotating blades of the compressor. Collectively one row of compressor vanes and one row of compressor blades make up a “stage” of the compressor. Similarly, the high pressure turbine includes annular rows of stationary nozzle vanes that direct the gases exiting the combustor into downstream, rotating blades of the turbine. Collectively one row of nozzle vanes and one row of turbine blades make up a “stage” of the turbine. Typically, both the compressor and turbine include a plurality of successive stages.
In order to allow for periodic inspection of the core parts of the engine (e.g., the compressor blades and the turbine blades), borescope ports are typically provided in the engine casings and/or frames. Such ports allow optical borescope instruments to be inserted into the core engine to enable a visual inspection of the engine to be performed without requiring disassembly of the engine components. However, once an instrument has been inserted into a borescope port, minimal information is typically available to an inspector regarding the actual position of the instrument within the engine, leading to errors in measurements and reducing the efficiency of performing the visual inspection.
Accordingly, a system and method for locating a probe relative to a gas turbine engine as such probe is being inserted within the engine would be welcomed within the technology.